Stator vane having reinforced braze joint

Abstract: A gas turbine engine stator vane 34 includes a groove 58 extending laterally in an outer surface 40 at one end 46 thereof between leading and trailing edges 42, 44 for receiving a brazing material 62. The vane 34 may be joined to an arcuate band 36 having a complementary groove 60 extending colinearly with the airfoil groove 58 to collectively define therebetween an enlarged crevice 66 for receiving the brazing material 62 to form a reinforcing pin 62a therein.

Description: The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein. BACKGROUND OF THE INVENTION In a gas turbine engine, a combustor discharges hot combustion gases downstream through a high pressure (HP) stationary or stator turbine nozzle which directs the flow between rotor blades of a high pressure turbine (HPT) for extracting energy therefrom. The HPT may have a second stage of rotor blades disposed downstream from the first stage with an additional, second stage turbine nozzle disposed therebetween for channeling the combustion gases from the first stage rotor blades to the second stage rotor blades. The HP first stage turbine nozzle and the second stage turbine nozzle each include a plurality of circumferentially spaced apart stator vanes or airfoils joined at their radially outer and inner ends to annular outer bands. The nozzles are typically made in arcuate segments with arcuate outer and inner band segments each having one or more vanes per segment. The segments are conventionally joined together to collectively form a complete 360° turbine nozzle. The HP stage one nozzle is typically supported in the engine at both its outer and inner bands for accommodating loads thereon including pressure forces from the combustion gases channeled between the vanes. However, the second stage nozzle is supported solely at its outer band since a conventional annular seal member is disposed between the first and second rotor stages preventing stationary support of the inner band as well. Accordingly, the vanes of the second stage nozzle are cantilevered from the outer band support which creates bending moments due to the combustion gases flowing between the vanes which must be suitably reacted or accommodated through the outer band. Since the HP stage one nozzle is supported at both its inner and outer bands, it may be relatively simply manufactured by brazing the vanes at their outer and inner ends to the respective outer and inner bands. However, braze joints have acceptable shear strength but undesirable bending strength. Since the HP nozzle is supported at its outer and inner bands, bending moments from the combustion gases are insignificant, whereas the bending moments in the second stage nozzle are significant since it is supported solely at its outer band, with the inner band thereof being unsupported. Accordingly, brazed turbine nozzles are typically not used where they cannot be supported at both their outer and inner bands which, therefore, requires alternate and typically more complex and expensive designs. For example, the vanes may be integrally cast in groups to their outer and inner bands to form integral nozzle segments which do not require brazing between the vanes and bands. Alternatively, a single vane may be integrally cast to outer and inner band segments to avoid brazing therebetween, with the band segments themselves being brazed together at the circumferential joints therebetween. However, these methods of manufacture are not suitable for conventional, high strength, single crystal vanes which are desirable for use in second stage turbine nozzles of improved gas turbine engines. Accordingly, a simpler and less expensive brazed turbine nozzle is desired for those stages wherein the nozzle is supported solely by its outer band, and which allows the use of single crystal vanes. DE-C-969599 shows a nozzle welded into a nozzlering with welds around individual nozzles. Summary of the Invention According to the invention, there is provided a gas turbine engine stator vane comprising a hollow airfoil having an outer surface extending laterally between a leading edge and a trailing edge and longitudinally between first and second opposite ends characterised in that said outer surface adjacent said first end has at least one groove therein extending at least in part laterally between said leading and trailing edges for receiving a brazing material. The vane may be joined to an arcuate band having an inner surface defining a support aperture being complementary in configuration with said airfoil as said first end and receiving therein said first end, said support aperture inner surface having at least one groove therein extending colinearly with and facing said airfoil groove to collectively define therebetween an enlarged crevice for receiving said brazing material. Brief Description of the Drawings The invention will now be described in greater detail, by way of example, with reference to the drawings in which: Figure 1 is an axial, partly sectional view of a portion of an exemplary gas turbine engine having a two-stage high pressure turbine with a turbine nozzle disposed therebetween in accordance with one embodiment of the present invention. Figure 2 is an exploded view of an exemplary segment of the second stage turbine nozzle illustrated in Figure 1 in accordance with one embodiment of the present invention. Figure 3 is a perspective view of the nozzle segment illustrated in Figure 2 in assembled form. Figure 4 is a transverse radial sectional view through the outer end of one of the vanes illustrated in Figure 3 adjacent to the outer band and taken along line 4-4 to illustrate a braze joint in accordance with one embodiment of the present invention. Figure 5 is a transverse radial sectional view through the outer end of one of the vanes illustrated in Figure 3 adjacent to the outer band and taken along line 4-4 to illustrate a braze joint in accordance with a second embodiment of the present invention. Figure 6 is a transverse radial sectional view through the outer end of one of the vanes illustrated in Figure 3 adjacent to the outer band and taken along line 4-4 to illustrate a braze joint in accordance with a third embodiment of the present invention. Figure 7 is a top view of a portion of the outer band and one of the vanes thereof illustrated in Figure 3 showing assembly of a first reinforcing strip therein. Figure 8 is a radial sectional view of the outer end of the vane and outer band illustrated in Figure 7 and taken along line 8-8. Figure 9 is a top view similar to Figure 7 illustrating the assembly of a second reinforcing strip between the vane and outer band. Figure 10 is a radial sectional view of the outer end of the vane and outer band illustrated in Figure 9 and taken along line 10-10. DESCRIPTION OF THE PREFERRED EMBODIMENT(S) Illustrated schematically in Figure 1 is a portion of an exemplary aircraft gas turbine engine 10 having in serial flow communication about a longitudinal or axial centerline axis a conventional axial, high pressure compressor (HPC) 12, combustor 14, high pressure (HP) turbine nozzle 16 (stage one), a high pressure turbine (HPT) stage one rotor 18, and an HPT stage two rotor 20 disposed downstream therefrom. During operation, the HPC 12 provides compressed air 22 to the combustor 14 wherein it is conventionally mixed with fuel and ignited to generate hot combustion gas 24. The combustion gas 24 is channeled through the HP nozzle 16 and through the rotor blades of the first stage rotor 18 which extract energy therefrom for driving the HPC 12. The combustion gas 24 also flows between the rotor blades of the second stage rotor 20 which extract additional energy therefrom. The HP nozzle 16 is conventional and includes a plurality of circumferentially spaced apart vanes joined between outer and inner bands which are separately supported by a conventional annular outer casing 26 and an annular inner support 28, respectively. In accordance with one embodiment of the present invention a second stage stationary or stator HPT turbine nozzle 30 is disposed between the first and second stage rotors 18, 20. The nozzle 30 is supported solely to the outer casing 26 since a conventional annular rotor seal member 32 is disposed radially below the nozzle 30 and axially between the first and second stage rotors 18, 20. More specifically, the nozzle 30 includes a plurality of circumferentially spaced apart stator vanes 34 as shown in Figures 1-3 which are joined to a radially outer or first band 36 and to a radially inner or second band 38. As shown in Figures 2 and 3, the nozzle 30 preferably includes a plurality of conventionally joined, circumferentially spaced apart, arcuate nozzle segments 30a having two vanes 34 per segment, for example, fixedly joined to respective arcuate portions of the outer and inner bands 36, 38 in accordance with the present invention. Although two vanes 34 per segment 30a are shown, one vane 34 per segment 30a or more than two vanes per segment 30a may be used as desired. As shown in Figure 2, each stator vane 34 is in the form of a hollow airfoil having an outer surface 40 extending laterally between a radially extending leading edge 42 and a radially extending trailing edge 44. The leading and trailing edges 42, 44 extend radially, or longitudinally relative to the vane 34, between a first or radially outer end 46 and an opposite second or radially inner end 48. The general configuration of the vane 34 is conventional with the outer surface 40 thereof defining an outwardly convex side and an opposite outwardly concave side extending laterally or axially between the leading and trailing edges 42, 44. The outer band 36 includes one or more first support apertures 50 defined by a radially extending inner surface 52 which is complementary in configuration with the vane 34 at the first end 46 thereof for receiving therein the first end 46. Similarly, the inner band 38 includes one or more second support apertures 54 defined by a radially extending inner surface 56 which is complementary in configuration with the vane 34 at the second end 48 thereof for receiving therein the second end 48. In accordance with one embodiment of the present invention, the vane outer surface 40 adjacent the first end 46 includes at least one U-shaped groove or slot 58 therein extending at least in part laterally between the leading and trailing edges 42, 44, and in the preferred embodiment illustrated in Figure 2, the groove 58 extends completely around the vane 34 along both sides thereof between the leading and trailing edged 42, 44. Similarly, the support aperture inner surface 52 includes at least one U-shaped groove or slot 60 positioned therein for extending colinearly with and facing the respective airfoil grooves 58 upon assembly. The band grooves 60 similarly extend completely around the inner surface 52 of the support aperture 50 in the preferred embodiment. As shown in Figure 2, each of the vanes 34 has its first end 46 inserted into its respective support aperture 50, with the second end 48 thereof being inserted into its respective support aperture 54 in the inner band 38 to form the four-piece assembly illustrated in Figure 3. The vanes 34 are conventionally fixtured to the outer and inner bands 36, 38 so that they may be conventionally brazed together using a conventional brazing material 62 in both the outer and inner bands 36, 38. The brazing material 62 is suitably melted and spread by capillary action between the outer and inner ends 46, 48 of the vanes 34 and the respective first and second support apertures 50, 54 which upon solidification rigidly joins the vanes 34 to the outer and inner bands 36, 38 to form the nozzle segment 30a. As illustrated in more particularity in Figure 4, the vane first end 46 is conventionally spaced apart from the support aperture inner surface 52 to define a gap 64 therebetween. The gap 64 has a predetermined thickness G for receiving the brazing material 62 which fills the gap 64 upon solidification after brazing for bonding together the vane first end 46 to the inner surface 52 of the first support aperture 50. As shown in Figures 2 and 4, the band grooves 60 extend colinearly with and face the respective airfoil grooves 58 to collectively define therebetween respective enlarged interstitial crevices 66 as illustrated in Figure 4 for receiving the brazing material 62 therein. The brazing material 62 is shown solidified in Figure 4 after the brazing operation for bonding together the vane first end 46 in the support aperture 50 of the outer band 36, and filling the crevice 66 to form an enlarged reinforcing pin 62a therein. In the exemplary embodiment illustrated in Figure 4, the airfoil groove 58 and the band groove 60 are generally semi-circular in transverse section so that the crevice 66 is generally circular in transverse section, with the corresponding pin 62a being substantially circular in transverse section. In the embodiment illustrated in Figure 4, a plurality of the airfoil grooves 58 and of the band grooves 60 are provided and disposed in radially spaced apart, or longitudinally spaced apart relative to the vane axis, pairs to define respective pluralities of the crevices 66 forming respective ones of the pins 62a therein. Alternatively, either a single crevice 66 and respective pin 62a may be formed or more than two crevices 66 and respective pins 62a may be formed as desired for suitably reinforcing the brazed joint between the vane outer end 46 and the outer band 36. Accordingly, the vane outer end 46 is fixedly joined to the outer band 36 through the first support aperture 50 by the brazing material 62 completely around the circumference thereof. The brazing material 62 forms the enlarged pins 62a in the respective crevices 66, which pins 62 also extend completely around the vane outer end 46 in this exemplary embodiment. The brazing material 62 is therefore intimately bonded to the vane outer surface 40 at its first end 46 and to the inner surface 52 of the first support aperture 50 for providing a brazed joint which is conventionally strong in shear. The reinforcing pins 62a provide additional strength to the brazed joint for accommodating the bending moment M illustrated in Figure 1 due to the cantilever support of the nozzle 30 at its outer band 36. More specifically, in the event of any separation between the brazing material 62 and the corresponding surfaces of the vane 34 and outer band 36 during operation in the engine 10, the pins 62a will remain as mechanical structures preventing disassembly of the vane 34 from the outer band 36. The pins 62a will therefore act as shear pins preventing separation of the components. Suitable conventional brazing material 62 may be used to provide an acceptably strong shear-pin type joint. And, any suitable filler or brazing material having a melting temperature at or below that of the vane 34 and band 36 may be used as desired for providing varying degrees of strength in the joint. For further increasing the strength of braze joint between the vane 34 and the outer band 36, for example when using lower strength brazing materials, an elongate reinforcing strip 68 as illustrated in Figure 5 may be disposed in the crevice 66 and along the airfoil groove 58 and the band groove 60 and within the pin 62a. As shown in Figure 5 the reinforcing strip 68 is rectangular in transverse section. As shown in the Figure 6 embodiment of the invention, the reinforcing strip designated 68a is circular in transverse section. The reinforcing strip 68, 68a may take any suitable configuration for reinforcing the pin 62a and the joint formed by the brazing material 62. Figures 7-10 illustrate an exemplary embodiment of the circular reinforcing strip 68a and its assembly between the vane 34 and the outer band 36. As shown in Figure 7, the reinforcing strip 68a may have any suitable length and in the exemplary embodiment illustrated, it is configured for extending substantially between the vane leading edge 42 to the vane trailing edge 44 on the convex side thereof. As shown in Figure 6 upon completion of the brazing process, the thickness of the braze 62 is nominally the gap thickness G which is preferably fairly uniform around the circumference of the vane 34 as is obtained in conventional practice and may be about 0.25 mm for example. The reinforcing strip 68a has a width W which is preferably greater than the gap thickness G. This is to ensure that the reinforcing strip 68a is as large as practical so that it is not readily removed during the assembly process and binds with the brazing material 62 substantially completely therearound to form the reinforcing pin 62a. In order to assemble the reinforcing strip 68a which is wider than the nominal gap thickness G, the vane 34 may be initially moved to the left as indicated by the arrow labeled L in Figures 7 and 8 so that the vane 34 touches the left side of the inner surface 52 of the first support aperture 50 and maximizes the gap at the right side of the vane 34 which has a magnitude of about 2G. The width W of the reinforcing strip 68a is suitably less than the maximum available gap thickness 2G so that it may be inserted downwardly into the gap 64 and into position within the crevice 66. A suitable temporary retention wire 70 may be inserted up to the crevice 66 from the bottom of the first support aperture 50 in order to temporarily hold the reinforcing strip 68a in position in the crevice 66 until the vane 34 is moved to the right as indicated by the arrow labeled R in Figures 9 and 10 as the retention wire 70 is removed. The reinforcing strip 68a on the right side of the vane 34 is then trapped in the crevice 66, and then a second one of the reinforcing strip 68a may be similarly inserted into the enlarged gap on the left side of the vane 34. The vane 34 may then be centered within the first support aperture 50 for obtaining the nominal gap G around its circumference and thusly trapping within the crevice 66 the two reinforcing strips 68a on both sides of the vane 34. The brazing material 62 is then conventionally applied to the gap 64 between the vane 34 and the outer band 36 and conventionally heated to flow by capillary action to completely fill the gap 64 around the circumference of the vane outer end 46, which after cooling forms a relatively rigid brazed joint. The braze joint of the vane outer end 46 and the outer band 36 has improved strength due to the so-formed shear pin 62a either alone or with the reinforcing strip 68, 68a therein. This allows the vanes 34 themselves to be conventional high strength, single crystal components, integrally joined to the conventional, non-single crystal outer and inner bands 36, 38. Since the inner band 38 as illustrated in Figure 1 is unsupported and is not subject to the bending moment M, it may be conventionally brazed to the vane inner ends 48 as shown in Figure 3 with the brazing material 62 filling the gap therebetween without the crevices 66 or reinforcing pin 62a formed therein. Of course, and if desired, the reinforcing pins 62a may also be formed between the vane inner end 48 and the inner band 38. The reinforcing strips 68, 68a may be any conventional metal such as Hastalloy X or other conventionally known nickel-based alloys. The reinforcing strips 68, 68a provide a mechanical locking feature of the braze joint and due to metallurgical bonding or interdiffusion with the brazing material 62 also provide additional strengthening of the reinforcing pin 62a, as well as limits the volume of the weaker brazing material 62a.

Claims: A gas turbine engine stator vane comprising a hollow airfoil (34) having an outer surface (40) extending laterally between a leading edge (42) and a trailing edge (44) and longitudinally between first and second opposite ends (46, 48) characterised in that said outer surface (40) adjacent said first end (46) has at least one groove (58) therein extending at least in part laterally between said leading and trailing edges (42, 44) for receiving a brazing material (62). A stator van according to claim 1 in combination with an arcuate first band (36), said first band (36) having an inner surface (52) defining a support aperture (50) being complementary in configuration with said airfoil (34) as said first end (46) and receiving therein said first end (46), said support aperture inner surface (52) having at least one groove (60) therein extending colinearly with and facing said airfoil groove (58) to collectively define therebetween an enlarged crevice (66) for receiving said brazing material 62. A combination according to claim 2 further including said brazing material (62) bonding together said airfoil first end (46) in said first band (36), and filling said crevice (66) to form a reinforcing pin (62a) therein. A combination according to claim 3 wherein said airfoil groove (58) and said band groove (60) are generally semi-circular in transverse section so that said crevice (66) is generally circular in transverse section. A combination according to claim 4 further including an elongate reinforcing strip (68, 68a) disposed in said crevice (66) along said airfoil groove (58) and said band groove (60), and within said pin (62a). A combination according to claim 5 wherein said airfoil first end (46) is spaced from said support aperture inner surface (52) to define a gap (64) therebetween having a predetermined thickness, and said strip (68, 68a) has a width greater that said gap thickness. A combination according to claim 6 wherein said strip (68) is rectangular in transverse section. A combination according to claim 6 wherein said strip (68a) is circular in transverse section. A combination according to claim 6 wherein said gap (64) is filled with said brazing material (62) bonding together said airfoil first end (46) to said support aperture inner surface (52). A combination according to claim 6 further comprising a plurality of said airfoil grooves (58) and said band grooves (60) disposed in longitudinally spaced apart pairs to define respective pluralities of said crevices (66) forming respective ones of said pins (62a) therein. A combination according to claim 6 further comprising an arcuate second band (38) having an inner surface (56) defining a second support aperture (54) being complementary in configuration with said airfoil (34) at said second (48) and receiving therein said second end (48), said airfoil end (48) being fixedly joined to said second band (38) by brazing material (62) to form a turbine nozzle segment (30a).

Inventors: CORREIA VICTOR HUGO SILVA, MCALLISTER KEVIN GREGORY, CORREIA, VICTOR HUGO SILVA, MCALLISTER, KEVIN GREGORY
Applicants: GEN ELECTRIC, GENERAL ELECTRIC COMPANY
Application year: 1994
Main classification: F01D9
IPC codes: F01D9